Circulation Control as a Roll Effector for Unmanned Combat Aerial Vehicles

his paper reports on a numerical investigation of the use of trailing-edge circulation control as a roll effector on a generic unmanned combat aerial vehicle: the DLR-F19 stability and control configuration. The Coanda effect induced by fluidic injections at the trailing edge of a wing is used to increase circulation and generate lift. Reynolds-averaged Navier–Stokes predictions are validated against wind-tunnel experiments conducted at the Georgia Institute of Technology and NASA’s basic aerodynamic research tunnel on an airfoil employing trailing-edge circulation control. Two turbulence models are used: the Wilcox k-ω model and Menter’s shear-stress transport, showing that the Wilcox k-ω model provides the best comparisons with the experimental data. Baseline data for the stability and control configuration with conventional control surfaces from wind-tunnel experiments done at the low speed wind-tunnel owned by the German- Dutch Wind Tunnels foundation (DNW-NWB) are used to ensure the correct flow features are being modeled for the flows encountered by this type of unmanned combat aerial vehicle and to provide a comparison for the performance of the circulation control devices. Modifications have been made to the DLR-F19, replacing the conventional control surfaces with trailing-edge circulation control of the same spanwise extent. This includes two configurations: one with a single slot, and one with three slots of equal width along the wing. The circulation control performs well at low angles of attack, producing a similar roll moment to the conventional control surfaces. Due to the flow separation at the high angles of attack, the circulation control is unable to generate a rolling moment. Finally, the flow topology is examined to understand the causes of the decrease in the performance

Detecting shape change: Characterizing the interaction between texture-defined and contour-defined borders

The human visual system's extreme sensitivity to subtle changes in shape can often be attributed to global pooling of local information. This has been shown for shapes described by paths of contiguous elements, but it was unknown whether this global pooling translated to shapes defined by texture-segmentation borders. Also, previous research suggests that texture and luminance cues-to-shape are integrated by the visual system for shape detection but it has not been established whether they combined for shape discrimination. Controlled shapes defined either by an explicit path of Gabors, texture-segmentation borders, or both of these cues were used. Results show that all stimuli used were globally processed. Thresholds for shapes defined by both cues matched predictions based on an independent-cue vector sum of individual thresholds. Thus, while local elements are integrated around the contour and are processed by global shape-detection mechanisms, integration did not occur across different shape-cues

ANN BASED ROM FOR THE PREDICTION OF UNSTEADY AEROELASTIC INSTABILITIES

Abstract. A Reduced-Order Model (ROM) for the prediction of aeroelastic instabilities is presented. The unsteady nonlinear aerodynamic system is characterised by an Artificial Neural Network (ANN) to a set of network weights. The system is trained on a time history of simultaneous forced oscillation of the normal modes as input and generalised forces as output. Network weights are then used to approximate the aerodynamic force in the structural equation of motion to obtain the structural response. Results from the 3D Goland wing are presented and compared against full order CFD. It is shown that the ROM can predict aeroelastic instabilities with reasonable accuracy at a cost of less than one typical unsteady aeroelastic computation

Nonlinear model order reduction for rapid gust loads analysis of flexible manoeuvring aircraft

This paper describes a systematic approach to the nonlinear model order reduction of free–flying flexible aircraft and the subsequent flight control system design. System nonlinearities arise due to large wing deformations and the coupling between flexible and rigid–body dynamics. The nonlinear flight dynamics equations are linearised and the approach uses information on the eigenspectrum of the resulting coupled system Jacobian matrix and projects it through a series expansion onto a small basis of eigenvectors representative of the full–order model dynamics. A very flexible aircraft representative of a high–altitude long–endurance (HALE) aircraft is implemented and the aeroelastic solver is verified against results from the literature. Furthermore, a very large flexible wing of high–aspect ratio is built and the flexibility effects on the flight dynamic response are investigated. The reduced–order model eigenvalue basis is identified and convergence studies are performed. Reduced–order models are generated and used for fast parametric worst–case gust searches of the full–order nonlinear flight dynamic response and are exploited in a robust control methodology for load alleviation. Finally, the performance of the controller is investigated on the nonlinear full–order model for various gust lengths

Numerical Simulation of Film Cooling in Hypersonic Flows

No abstract available

Analysis of the Boeing 747-100 using CEASIOM

One of the requirements for the SimSAC project was to use existing aircraft to act as benchmarks for comparison with CEASIOM generated models. Within this paper, results are given for one of these examples, the Boeing 747-100. This aircraft was selected because a complete dataset exists in the open domain, which can be used to validate SimSAC generated data. The purpose of this paper is to both give confidence in, and to demonstrate the capabilities of, the CEASIOM environment when used for preliminary aircraft and control system design. CEASIOM is the result of the integration of a set of sophisticated tools by the European Union funded, Framework 6 SimSAC program. The first part of this paper presents a comparison of the aerodynamic results for each of the solvers available within CEASIOM together with data from the 747-100 model published by NASA. The resulting nonlinear model is then trimmed and analysed using the Flight Control System Designer Toolkit (FCSDT) module. In the final section of the paper a state-feedback controller is designed within CEASIOM in order to modify the longitudinal dynamics of the aircraft. The open and closed loop models are subsequently evaluated with selected failed aerodynamic surfaces and for the case of a single failed engine. Through these results, the CEASIOM software suite is shown to be able to generate excellent quality adaptive-fidelity aerodynamic data. This data is contained within a full nonlinear aircraft model to which linear analysis and control system design can be easily applied

Aeroelastic System Identification using Transonic CFD data for a 3D Wing

peer reviewedThis paper is part of a study investigating the prediction of aeroelastic behaviour subjected to non-linear aerodynamic forces. Of interest here is whether the sub-critical vibration behaviour of the aeroelastic model gives any information about the onset of the LCO. It would be useful to be able to use system identification methods to estimate aeroelastic models that characterise the LCO. Such a methodology would be very useful, not only for analysis with coupled CFD/FE models, but also during flight flutter testing. In this paper, the responses to initial inputs on the Goland Wing [9] CFD/FE model at different flight speeds are analysed to determine the extent of the non-linearity below the critical onset of LCO. Analysis is also performed using a linear identification model

Aeroelastic system identification using transonic CFD data for a wing/store configuration

This paper is part of a study investigating the prediction of the aeroelastic behaviour of aircraft subjected to non-linear aerodynamic forces. The main objective of the work is the characterization of the dynamic response of aeroelastic models resulting from coupled Computational Fluid Dynamic and Finite Element calculations. Of interest here is the identification of the flight condition at which the response bifurcates to limited or divergent amplitude self-sustained oscillations without carrying out a comprehensive set of full, computationally expensive, time-marching calculations. The model treated in this work is a three-dimensional wing in a transonic flowfield. Short datasets of pre-bifurcation behaviour are analysed to determine the system’s stability and degree of non-linearity. It is found that the calculated responses on the run-up to a transonic Limit Cycle Oscillation show little or no evidence of non-linearity. The non-linearity appears abruptly at the bifurcation flight condition. The variation of the local Mach number over the wing’s surface in the steady-state case is used to demonstrate that the non-linearity is due to a shock wave that can move along the surface. At Mach numbers where this is not possible the system behaves in a linear manner and its stability can be analysed using linear methods